Abstract: An assembly is provided for a turbine engine. This assembly includes a housing, a combustor and a steam injector. The housing includes an air plenum. The combustor is disposed within the air plenum. The combustor includes a combustor wall and a combustion chamber. The combustor wall is disposed between the combustion chamber and the air plenum. The combustor wall includes a quench aperture that extends through the combustor wall from the air plenum to the combustion chamber. The steam injector projects partially into or through the quench aperture.

Abstract: A gas turbine engine fuel system includes a fuel delivery system, an oil cooling subsystem, and a fuel conditioning subsystem. The fuel delivery subsystem delivers fuel to a gas turbine engine, and the oil cooling subsystem receives heated oil from the gas turbine engine. The fuel conditioning subsystem includes a fuel/oil cooler that is in fluid communication with the fuel delivery subsystem to receive the fuel and is in fluid communication with the oil cooling subsystem to receive the heated oil, the fuel/oil cooler configured to transfer heat from the heated oil to the fuel. A thermal coking sensor is in fluid communication with the fuel and is configured to generate a signal in response to interacting with the fuel. A controller is configured to determine a coking temperature indicating an onset of coking in the fuel based on the signal.

Abstract: A gas turbine engine fuel system includes a fuel delivery system, an oil cooling subsystem, a fuel delivery conditioning system, and one or more optical coking sensors. The fuel delivery subsystem delivers fuel to a gas turbine engine and the oil cooling subsystem receives heated oil from the gas turbine engine. The fuel delivery conditioning system includes a fuel/oil cooler that is in fluid communication with the fuel delivery subsystem to receive the fuel and is in fluid communication with the oil cooling subsystem to receive the heated oil. The fuel/oil cooler transfers heat from the oil to the fuel to increase a temperature of the fuel, and the optical coking sensors output a signal indicating a thermal stress of the fuel that occurs with a changing temperature of the fuel. A controller receives the voltage signal determines an onset of coking in the fuel based on the signal.

Abstract: An operating method is provided for an aircraft powerplant. During this method, fuel is directed in a first fuel circuit from a fuel source towards a fuel injector in a combustion section of the aircraft powerplant. A fuel system for the aircraft powerplant includes the fuel source, the first fuel circuit and the fuel injector. The fuel system is configured to fluidly couple the fuel source to the fuel injector through the first fuel circuit. A phase parameter of the fuel within the first fuel circuit is monitored at a monitoring location between the fuel source and the fuel injector. The fuel system is operated based on the phase parameter. The phase parameter may be indicative of the fuel at the monitoring location being in a gaseous phase. The phase parameter may be indicative of the fuel at the monitoring location being in a liquid phase or a mixed phase.

Abstract: A gas turbine engine fuel system includes a fuel delivery system, an oil cooling subsystem, a fuel delivery conditioning system, and one or more optical coking sensors. The fuel delivery subsystem delivers fuel to a gas turbine engine and the oil cooling subsystem receives heated oil from the gas turbine engine. The fuel delivery conditioning system includes a fuel/oil cooler that is in fluid communication with the fuel delivery subsystem to receive the fuel and is in fluid communication with the oil cooling subsystem to receive the heated oil. The fuel/oil cooler transfers heat from the oil to the fuel to increase a temperature of the fuel, and the optical coking sensors output a signal indicating a thermal stress of the fuel that occurs with a changing temperature of the fuel. A controller receives the voltage signal determines an onset of coking in the fuel based on the signal.

Abstract: A gas turbine engine fuel system includes a fuel delivery system, an oil cooling subsystem, and a fuel conditioning subsystem. The fuel delivery subsystem delivers fuel to a gas turbine engine, and the oil cooling subsystem receives heated oil from the gas turbine engine. The fuel conditioning subsystem includes a fuel/oil cooler that is in fluid communication with the fuel delivery subsystem to receive the fuel and is in fluid communication with the oil cooling subsystem to receive the heated oil, the fuel/oil cooler configured to transfer heat from the heated oil to the fuel. A thermal coking sensor is in fluid communication with the fuel and is configured to generate a signal in response to interacting with the fuel. A controller is configured to determine a coking temperature indicating an onset of coking in the fuel based on the signal.

Abstract: A gas turbine engine fuel system includes a fuel delivery system, an oil cooling subsystem, and a fuel conditioning subsystem. The fuel delivery subsystem delivers fuel to a gas turbine engine, and the oil cooling subsystem receives heated oil from the gas turbine engine. The fuel conditioning subsystem includes a fuel/oil cooler that is in fluid communication with the fuel delivery subsystem to receive the fuel and is in fluid communication with the oil cooling subsystem to receive the heated oil, the fuel/oil cooler configured to transfer heat from the heated oil to the fuel. A thermal coking sensor is in fluid communication with the fuel and is configured to generate a signal in response to interacting with the fuel. A controller is configured to determine a coking temperature indicating an onset of coking in the fuel based on the signal.

Abstract: An assembly is provided for a turbine engine. This assembly includes a turbine vane array and a cooling system. The turbine vane array includes an inner platform, an outer platform and a plurality of turbine vanes extending between and connected to the inner platform and the outer platform. The turbine vanes include a first turbine vane, and the first turbine vane includes a first passage and a second passage. The cooling system is configured to direct a first fluid into the first passage and a second fluid into the second passage. The first fluid includes air and steam during a first mode. The second fluid includes the air without the steam during the first mode.

Abstract: A combustor including: a combustor case defining a plurality of case apertures; a liner within the combustor case defining a combustion zone and liner apertures through which an airflow flows into the combustion zone; a fuel injector having a fuel channel extending through a first case aperture and the liner, and the fuel channel has a nozzle at the combustion zone through which fuel is injected; an igniter for igniting the combustible mixture of fuel and airflow and providing a flame at the nozzle; and a flame sensor including: a radio frequency transponder, comprising a transmitter-receiver pair, located exterior to the combustor case; a horn antenna disposed in the fuel nozzle, and a tubular waveguide extending from the radio frequency transponder to the horn via one of the plurality of case apertures, wherein the flame sensor is configured to perform flame and flow field diagnostics.

Abstract: An assembly for a combustor for a gas turbine engine, including: a plurality of heat shield panels attached to at least one combustor liner, each one of the plurality of heat shield panels including a panel portion and a first forward rail and a second rearward rail each extending from an outer surface of the panel portion, the panel portion including an inner surface opposite of the outer surface, the panel portion also includes a forward end and a rearward end, the forward end of the panel portion is axially forward of the rearward end of an adjacent heat shield panel of the plurality of heat shield panels such that a gap is defined between the outer surface of the panel portion of one heat shield panel of the plurality of heat shield panels and an inner surface of the panel portion of an adjacent heat shield panel of the plurality of heat shield panels, the panel portion also includes a plurality of apertures extending from the outer surface to the inner surface; and a plurality of cooling pins extending u

Abstract: During a method, an aircraft powerplant is provided. The aircraft powerplant includes an injector and a fuel system configured to deliver fuel to the injector. The fuel system includes a first tank, a first circuit and a second circuit arranged in parallel with the first circuit between the first tank and the injector. The second circuit includes a second tank. The fuel is directed from the first tank to the injector during a first mode through the first circuit and independent of the second circuit. The fuel is directed from the second tank to the injector during a second mode through the second circuit and independent of the first circuit.

Abstract: A fuel injector assembly for a gas turbine engine includes a swirler and a fuel nozzle. The swirler has an outer wall having a SOW inner wall surface, an inner wall having a SIW inner wall surface and a SIW outer wall surface, an outer passage, and an inner passage. The SIW inner wall surface includes first and second sections. The first section extends between a first airflow inlet and the second section, and the second section extends between the first section and an inner wall distal end. The first section tapers radially inward, and the second section extends parallel the center axis. The outer wall circumscribes the inner wall and extends axially along the center axis to a distal outer wall end, and the distal end wall is axially recessed within the swirler from the distal outer wall end. The fuel nozzle projects into the inner passage.

Abstract: A combustion assembly for a gas turbine engine includes a combustor, a monochromator, and a photodetector assembly. The combustor forms a combustion chamber. The monochromator is disposed outside the combustion chamber. The monochromator is configured to receive an optical input from the combustion chamber and direct an optical output. The optical input has a range of light wavelengths. The optical output has a subset of the range of light wavelengths. The photodetector assembly is disposed outside the combustion chamber. The photodetector assembly is configured to receive the optical output from the monochromator and generate an output signal representative of one or more optical characteristics of the optical output.

Abstract: An assembly for a combustor for a gas turbine engine, including: a plurality of heat shield panels attached to at least one combustor liner, each one of the plurality of heat shield panels including a panel portion and a first forward rail and a second rearward rail each extending from an outer surface of the panel portion, the panel portion including an inner surface opposite of the outer surface, the panel portion also includes a forward end and a rearward end, the forward end of the panel portion is axially forward of the rearward end of an adjacent heat shield panel of the plurality of heat shield panels such that a gap is defined between the outer surface of the panel portion of one heat shield panel of the plurality of heat shield panels and an inner surface of the panel portion of an adjacent heat shield panel of the plurality of heat shield panels, the panel portion also includes a plurality of apertures extending from the outer surface to the inner surface; and a plurality of cooling pins extending u

Abstract: An assembly is provided for a turbine engine. This assembly includes a housing, a combustor and a steam injector. The housing includes an air plenum. The combustor is disposed within the air plenum. The combustor includes a combustor wall and a combustion chamber. The combustor wall is disposed between the combustion chamber and the air plenum. The combustor wall includes a quench aperture that extends through the combustor wall from the air plenum to the combustion chamber. The steam injector projects partially into or through the quench aperture.

Abstract: An assembly is provided for a turbine engine. This assembly includes a housing, a combustor and a fuel injector assembly. The housing includes a case and a diffuser plenum radially within the case. The case extends axially along and circumferentially about a centerline. The combustor is disposed within the diffuser plenum. The combustor includes a bulkhead, a hood, a combustion chamber and a hood cavity. The bulkhead is axially between the combustion chamber and the hood cavity. The hood is between the hood cavity and the diffuser plenum. The fuel injector assembly projects radially into the diffuser plenum from the case. The fuel injector assembly includes a fuel nozzle and a steam nozzle. The fuel nozzle is configured to inject fuel into the combustion chamber. The steam nozzle is configured to inject steam into the hood cavity.

Abstract: An assembly is provided for a turbine engine. This assembly includes a housing, a combustor and a steam injector. The housing includes an air plenum. The combustor is disposed within the air plenum. The combustor includes a combustor wall and a combustion chamber. The combustor wall is disposed between the combustion chamber and the air plenum. The combustor wall includes a quench aperture that extends through the combustor wall from the air plenum to the combustion chamber. The steam injector projects partially into or through the quench aperture.

Abstract: A fuel injector assembly for a gas turbine engine includes a swirler and a fuel nozzle. The swirler has an outer wall having a SOW inner wall surface, an inner wall having a SIW inner wall surface and a SIW outer wall surface, an outer passage, and an inner passage. The SIW inner wall surface includes first and second sections. The first section extends between a first airflow inlet and the second section, and the second section extends between the first section and an inner wall distal end. The first section tapers radially inward, and the second section extends parallel the center axis. The outer wall circumscribes the inner wall and extends axially along the center axis to a distal outer wall end, and the distal end wall is axially recessed within the swirler from the distal outer wall end. The fuel nozzle projects into the inner passage.

Abstract: An aircraft propulsion system is provided that include a propulsor rotor, a turbine engine core and a steam system. The turbine engine core is configured to power the propulsor rotor. The turbine engine core includes a core flowpath, a compressor section, a combustor section and a turbine section. The core flowpath extends through the compressor section, the combustor section and the turbine section from a flowpath inlet to a flowpath exhaust. The combustor section includes a diffuser plenum and a combustor disposed within the diffuser plenum. The steam system is configured to introduce steam into the core flowpath upstream of the diffuser plenum.

Abstract: An assembly is provided for a turbine engine. This assembly includes a housing, a combustor and a fuel injector assembly. The housing includes a case and a diffuser plenum radially within the case. The case extends axially along and circumferentially about a centerline. The combustor is disposed within the diffuser plenum. The combustor includes a bulkhead, a hood, a combustion chamber and a hood cavity. The bulkhead is axially between the combustion chamber and the hood cavity. The hood is between the hood cavity and the diffuser plenum. The fuel injector assembly projects radially into the diffuser plenum from the case. The fuel injector assembly includes a fuel nozzle and a steam nozzle. The fuel nozzle is configured to inject fuel into the combustion chamber. The steam nozzle is configured to inject steam into the hood cavity.